This application relates generally to gas turbine engine rotor blades and, more particularly, to methods and apparatus for reducing rotor blade tip temperatures.
Gas turbine engine rotor blades typically include airfoils having leading and trailing edges, a pressure side, and a suction side. The pressure and suction sides connect at the airfoil leading and trailing edges, and span radially between the airfoil root and the tip, such that a cooling cavity is defined therein. To facilitate reducing combustion gas leakage between the airfoil tips and stationary stator components, the airfoils include a tip region that extends radially outward from the airfoil tip.
The airfoil tip regions include a first tip wall extending from the airfoil leading edge to the trailing edge, and a second tip wall also extending from the airfoil leading edge to connect with the first tip wall at the airfoil trailing edge. The tip region prevents damage to the airfoil if the rotor blade rubs against the stator components.
During operation, combustion gases impacting the rotating rotor blades transfer heat into the blade airfoils and tip regions. Over time, continued operation in higher temperatures may cause the airfoil tip regions to thermally fatigue. To facilitate reducing operating temperatures of the airfoil tip regions, at least some known rotor blades include slots within the tip walls to permit combustion gases at a lower temperature to flow through the tip regions.
To facilitate minimizing thermal fatigue to the rotor blade tips, at least some known rotor blades employ a plurality of tip openings which enable cooling air to discharge from the cooling cavity for cooling the rotor blade tips. After assembly in the rotor, the rotor blades are typically ground as an assembly to a specified diameter for the rotor. During the assembly grind, and during engine operation, the tip openings may become smeared shut, thus decreasing an amount of cooling air that may be discharged from the cooling cavity for cooling the rotor blade tips. Rotor operation with rotor blade tip openings that have been smeared shut may increase the operating temperature of the rotor blade tips, which may result in an increased amount of thermal creep, plastic deformation, and oxidation formation within the trailing edge tip region. Over time, continued operation with rotor blade tip openings that have been smeared shut may decrease the useful life of the rotor assembly.
In one aspect of the invention, a method for fabricating a rotor blade for a gas turbine engine to facilitate reducing operating temperatures of a tip portion of the rotor blade is provided. The method comprises forming an airfoil including a first sidewall and a second sidewall connected at a leading edge and a trailing edge to define a cavity therein, wherein the first and second sidewalls extend radially between a rotor blade root and a rotor blade tip, and forming a first tip wall extending from the rotor blade tip plate along the first sidewall, such that at least a portion of the first tip wall is at least partially recessed with respect to the rotor blade first sidewall to define a tip shelf that extends from the airfoil trailing edge towards the airfoil leading edge. The method also comprises forming a second tip wall extending from the rotor blade tip plate along the second sidewall such that the second tip wall connects with the first tip wall at the rotor blade trailing edge.
In a further aspect, an airfoil for a gas turbine engine is provided. The airfoil includes a leading edge, a trailing edge, a tip plate, a first sidewall, a second sidewall, a first tip wall, and a second tip wall. The first tip wall extends in radial span between an airfoil root and the tip plate. The second sidewall is connected to the first sidewall at the leading edge and the trailing edge to define a cavity therein. The second sidewall extends in radial span between the airfoil root and the tip plate. The first tip wall extends radially outward from the tip plate along the first sidewall. The second tip wall extends radially outward from the tip plate along the second sidewall. The first tip wall is connected to the second tip wall at the leading edge. The first tip wall is at least partially recessed with respect to the rotor blade first sidewall to define a tip shelf that extends from the airfoil trailing edge towards the airfoil leading edge. The tip shelf includes at least one opening extending therethrough to the cavity.
In another aspect of the invention, a gas turbine engine including a plurality of rotor blades is provided. Each rotor blade includes an airfoil including a leading edge, a trailing edge, a first sidewall, a second sidewall, a first tip wall, and a second tip wall. The airfoil first and second sidewalls are connected axially at the leading and trailing edges to define a cavity within the airfoil. The first and second sidewalls extend radially from a blade root to the tip plate. The first tip wall extends radially outward from the tip plate along the first sidewall. The second tip wall extends radially outward from the tip plate along the second sidewall. The first tip wall is at least partially recessed with respect to the rotor blade first sidewall to define a tip shelf that extends from the airfoil trailing edge towards the airfoil leading edge. The tip shelf includes at least one opening extending therethrough to the airfoil cavity.